The exhaust plume of gas turbine engines, e.g., military aircraft engines, is a source of high infrared energy which may be used for targeting/tracking purposes by heat seeking missiles and/or various forms of infrared imaging systems, e.g., Night Vision Goggles (NVGs). With respect to the former, heat seeking missiles obtain directional cues from the infrared energy, which directional cues are more precise depending upon the intensity, e.g., temperature, and total energy emitted by the exhaust plume. Regarding the latter, infrared imaging systems amplify the infrared energy by a factor of about 10,000 and may be used by hostile forces for early detection and/or targeting of surface-to-air/air-to-air missiles. Accordingly, it is highly desirable to reduce the infrared emissions (also referred to as the IR signature) of such engines to a practical minimum to avoid detection, acquisition and/or tracking by enemy threats/forces.
Various infrared suppression systems have been designed and fielded which effect mixing of low temperature ambient air with the high temperature engine exhaust to reduce the IR signature radiated therefrom. Generally, it is the principle objectives of such systems to: (a) reduce the infrared energy below a threshold level (e.g., a level capable of being sensed by the perceived threat), (b) maintain engine performance, and, (c) minimize the weight penalties associated therewith. Secondary objectives may include: (i) minimizing system or configuration complexity to reduce fabrication costs, (ii) minimizing the external aerodynamic drag produced by such IR suppressors, and/or (iii) suppressing the acoustic emissions emanating therefrom which may also be a source of detection.
Hurley U.S. Pat. No. 4,018,046 discloses an InfraRed (IR) suppressor for reducing the infrared radiation emitted from the exhaust plume and/or hot metal parts of a gas turbine engine. The IR suppressor includes an ejector vane assembly which is adapted for receiving a primary flow of engine exhaust and which is operative to entrain a first flow of cooling air into the primary exhaust flow. The primary flow of gases diffused from and entrained by the ejector vane assembly are fed into an arcuate duct which, due to its spatial position relative to the ejector vane assembly, produces an annular space for entraining a second flow of cooling air. The first and second flows are admixed with the primary exhaust flow to effect heat transfer therebetween and, consequently, to suppress the IR signature emitted from the exhaust plume. Additionally, the curved shape of the arcuate duct and the use of turning vanes disposed internally thereof, serve to prevent direct line-of-sight viewing into the IR suppressor. That is, the shape and/or occluding characteristics of the turning vanes prevents viewing of radiant energy emitted by hot internal components such as from the ejector vane assembly or from the turbine section of the engine.
The ejector vane assembly described therein includes a plurality of V-shaped ejector vanes which entrain cooling air at a plurality of angularly spaced positions while additionally serving to introduce air toward the center or "core" of the engine exhaust. More specifically, each of the ejector vanes forms a V-shaped channel which is open to ambient air at one end thereof and is oriented to direct flow radially toward the core of the primary exhaust flow. The ejector vane assembly, furthermore, clusters the ejector vanes as close as practicable to the core of the primary exhaust flow and, furthermore, includes a means for adjusting the flow area through the ejector vanes to ameliorate engine performance. That is, the adjustment means effects collapse of each ejector vane so as to increase the flow area and reduce the back-pressure on the engine.
While the ejector vane assembly improves mixing, i.e., reduces the IR signature, by directing cooling air into a core region of the engine exhaust, it will be appreciated that the ejector vane assembly requires a large number of individual components. As such, the multiplicity of components, each being a potential source of repair and maintenance, reduces the overall reliability of the IR suppressor system. Furthermore, while the adjustment means attempts to reduce flow restriction, the anticipated power loss produced by the ejector vane assembly is on the order of between 3% to 6%. In view of the fact that even small variations in efficiency, i.e., on the order of 1/2% to 1%, have significant impact on engine performance, such power loss may be viewed as intolerable for certain aircraft and/or aircraft missions wherein maximum gross weight, flight speed, and range are critical design parameters.
Furthermore, while the turning vanes serve to redirect the exhaust flow and prevent direct line-of-sight viewing internally of the arcuate duct, impingement of the high temperature exhaust gases on the surfaces thereof can cause the turning vanes to become a significant source of infrared radiation. That is, depending upon the temperature reduction effected by the first stage of the suppressor, i.e., the ejector vane assembly, the temperature of the turning vanes can significantly contribute to the IR signature of the suppressor.
Miller U.S. Pat. No. 4,312,480 discloses a gas diffusion and radiation shielding apparatus for a turbine engine which provides multiple stages for entraining cooling air into a primary flow of high temperature engine exhaust, and a plurality of channels for segmenting the primary flow into smaller, more efficient, mixing streams. Furthermore, the channels include sidewalls which are functionally equivalent to turning vanes for redirecting the primary flow and preventing direct line of sight viewing to the interior of the apparatus/engine. Moreover, the sidewalls define internal conduits for entraining a cooling flow of air to reduce the surface temperature of the sidewall and the IR signature emitted thereby. More specifically, each internal conduit includes an inlet disposed along the upper and lower edges of the channel and an outlet disposed along a leading edge which opposes the primary flow of exhaust gases. In operation, the high-velocity low-pressure primary flow draws a "reverse flow" of cooling air through the conduit which is diffused via the leading edge outlet. As such, the cooling air reduces the surface temperature of the sidewalls and, consequently, the IR signature.
While IR radiation emitted by the channels, and more specifically, the sidewalls, is reduced via the reverse flow of cooling air, the reduction in IR signature is marginally beneficial, especially when compared to the cost of fabricating such internal conduits. Furthermore, by directing cooling air outwardly from the leading edge of the sidewall, back-pressure is produced which degrades engine performance.
Presz et al. U.S. Pat. Nos. 4,835,961 and 4,830,315 describe nozzle configurations for mixing/pumping fluid. More specifically, the Presz '961 and '315 patents describe single and multi-stage mixer/ejector nozzles, respectively, each having a plurality of adjoined lobes formed at an outlet end thereof. The adjoined lobes define a primary flow trough for channeling a primary flow of high velocity/energy fluid, such as the exhaust of a gas turbine engine, and a secondary flow trough for channeling a secondary flow of low velocity/energy fluid such as ambient air. The primary and secondary flow troughs are alternately disposed about the periphery of the nozzle such that thin sheets of high energy fluid flow from the trough outlets, transfer kinetic energy to the low energy fluid, and entrain the secondary flow into the primary flow. Aside from simple viscous or shear mixing, the adjoined lobes produce axial vortices which rapidly admix the primary and secondary flows. As such, the adjoined lobes described in the Presz '961 and '315 patents are known to have utility on aircraft engines to both increase thrust and for suppressing the thermal energy radiated from the engine.
While the adjoined lobes serve to rapidly admix the primary and secondary flows to reduce the IR signature, the nozzle configurations disclosed therein fail to fully mix the flows, particularly in the core region of the primary exhaust flow. That is, in all of the embodiments described therein, the lobes entrain the secondary cooling flow into a perimeter region of the flow such that a central core or band of high temperature exhaust remains as the flow exits the nozzle. While the core of high temperature flow will eventually admix downstream, it will be appreciated that the degree of IR suppression is compromised until such energy is diffused.
Aside from the IR emissions radiated by the engine, yet another source of detection includes the observable "radar signature" of the aircraft. The radar signature of an aircraft may be defined as the level of electromagnetic energy returned, i.e., reflected back, to a scanning radar when the aircraft is in the detection range or interrogation field of the radar. Depending upon the radar signature of the aircraft, the scanning radar may provide early detection so that appropriate countermeasures may be taken, or, if the radar signature is below the detection level of the scanning radar, may be entirely ineffective. Accordingly, for aircraft conducting covert operations, it is highly desirable to reduce the radar signature of the aircraft to obviate an enemy's ability to conduct countermeasures.
While many techniques are available to reduce the radar signature of an aircraft, the most common methods include select shaping of aircraft structure to direct the impinging radar away from its source, utilizing low-dielectric materials to make aircraft structure "invisible" to radar, and/or employing radar absorbent materials to attenuate to radar energy. As a general rule, it is desirable to maintain smooth exterior surfaces having relatively constant electrical properties for reflecting and/or attenuating the radar energy.
Insofar as IR suppressors typically require a complex geometric profile which cannot be readily "shaped" to reduce radar signature, it is common practice to apply radar absorbent materials to the exposed surfaces of such suppressors to effect signature reduction. Other approaches may involve select placement of the IR suppressor outlet which placement prevents radar line-of-sight viewing relative to an anticipated radar azimuth angle. For example, by placing the outlet upward, the aircraft fuselage may effectively occlude the suppressor outlet. Notwithstanding the method employed, it will be appreciated that the impact on radar signature is exacerbated on aircraft requiring multiple IR suppression apparatus.